Cooling air for gas turbine engine with thermally isolated cooling air delivery

ABSTRACT

A gas turbine engine includes a plurality of rotating components housed within a compressor section and a turbine section. A first tap is connected to the compressor section and configured to deliver air at a first pressure. A heat exchanger is connected downstream of the first tap. A flowpath is defined between a rotating surface and a non-rotating surface. The flowpath is connected downstream of the heat exchanger and is configured to deliver air to at least one of the plurality of rotating components. At least a portion of the non-rotating surface and the rotating surface includes a base metal. An insulation material is disposed on a surface along the flowpath.

CROSS-REFERENCE TO RELATED APPLICATION

This application is a continuation of U.S. patent application Ser. No.16/375,892 filed on Apr. 5, 2019, which claims priority to U.S.Provisional Application No. 62/653,647 filed on Apr. 6, 2018.

BACKGROUND

This application relates to the supply of cooled cooling air to variousrotating components in a gas turbine engine and wherein a cooling airpath is thermally isolated from hotter sections of the gas turbineengine.

Gas turbine engines are known and typically include a fan delivering airinto a bypass duct as bypass air and into a compressor in a core engine.The air in the compressor is compressed and delivered into a combustorwhere it is mixed with fuel and ignited. Products of this combustionpass downstream over turbine rotors driving them to rotate.

As can be appreciated, many of the components in a gas turbine enginesee very high temperatures. As an example, the turbine section and, inparticular, its early stages see hot products of combustion. Inaddition, the compressor and, in particular, its downstream most stagesalso see very high temperatures. This is particularly true as thepressures develop in the compressor sections are increasing.

Thus, it is known to supply cooling air to various rotating componentssuch as in the turbine section and/or compressor section.

SUMMARY

In a featured embodiment, a gas turbine engine includes a plurality ofrotating components housed within a compressor section and a turbinesection. A first tap is connected to the compressor section andconfigured to deliver air at a first pressure. A heat exchanger isconnected downstream of the first tap. A flowpath is defined between arotating surface and a non-rotating surface. The flowpath is connecteddownstream of the heat exchanger and is configured to deliver air to atleast one of the plurality of rotating components. At least a portion ofthe non-rotating surface and the rotating surface includes a base metal.An insulation material is disposed on a surface along the flowpath.

In another embodiment according to the previous embodiment, there is adownstream most location in a high pressure compressor within thecompressor section and the first tap is at an upstream location relativeto the downstream most location.

In another embodiment according to any of the previous embodiments,there is a high pressure compressor in the compressor section with adownstream most location and the first tap is at a location where airwill have passed downstream of the downstream most location.

In another embodiment according to any of the previous embodiments, theat least one rotating component includes at least a downstream mostportion of a high pressure compressor within the compressor section.

In another embodiment according to any of the previous embodiments, theat least one rotating component includes an upstream most blade and vanein a high pressure turbine which is part of the turbine section.

In another embodiment according to any of the previous embodiments, thematerial is provided outwardly of the base metal on at least a portionof both the rotating surface and the non-rotating surface.

In another embodiment according to any of the previous embodiments,there is a combustor radially outward of the non-rotating surface. Achamber is intermediate the combustor and the non-rotating surface isconnected to receive compressed air downstream of a downstream mostlocation in the compressor section.

In another embodiment according to any of the previous embodiments, therotating surface is an outer surface of a shaft connecting a highpressure turbine rotor in the turbine section to a high pressurecompressor rotor in the compressor section.

In another embodiment according to any of the previous embodiments, theinsulation material on the rotating surface is a coating.

In another embodiment according to any of the previous embodiments, thecoating includes an outer ceramic topcoat facing the insulation materialon the non-rotating surface.

In another embodiment according to any of the previous embodiments,there is a metallic bond coat intermediate said ceramic topcoat and theunderlying base metal in the rotating surface.

In another embodiment according to any of the previous embodiments,there is a thermally-grown oxide coating intermediate the metallic bondcoat and the ceramic topcoat.

In another embodiment according to any of the previous embodiments, theinsulation material on the non-rotating surface includes a ceramic fiberblanket.

In another embodiment according to any of the previous embodiments, thebase metal is radially inward of the ceramic fiber blanket and an outerwall of the non-rotating surface is attached on an opposed radial sideof the ceramic fiber blanket relative to the base metal.

In another embodiment according to any of the previous embodiments, theinsulation material on the non-rotating component includes a ceramicfiber blanket. The base metal is radially inward of the ceramic fiberblanket and an outer wall of the non-rotating structure is attached onan opposed radial side of the ceramic fiber blanket relative to the basemetal.

In another embodiment according to any of the previous embodiments,fluid conduits are connected to a location downstream of the heatexchanger, to communicate air downstream of the heat exchanger into theflow path, and at least some of the fluid conduits being provided withinsulation.

In another embodiment according to any of the previous embodiments, theceramic fiber blanket is formed of bulk fibers.

In another embodiment according to any of the previous embodiments,ceramic fiber blanket is formed with a alumina-silica fibers.

In another embodiment according to any of the previous embodiments,fluid conduits are connected to a location downstream of the heatexchanger, to communicate air downstream of the heat exchanger into theflow path, and at least some of the fluid conduits being provided withinsulation.

In another embodiment according to any of the previous embodiments,fluid conduits are connected to a location downstream of the heatexchanger, to communicate air downstream of the heat exchanger into theflow path, and at least some of the fluid conduits being provided withinsulation.

These and other features may be best understood from the followingdrawings and specification.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 schematically shows a gas turbine engine.

FIG. 2 shows a first cooling air supply system.

FIG. 3 shows an alternative system.

FIG. 4 shows a detail of the FIG. 2 or 3 system.

DETAILED DESCRIPTION

FIG. 1 schematically illustrates a gas turbine engine 20. The gasturbine engine 20 is disclosed herein as a two-spool turbofan thatgenerally incorporates a fan section 22, a compressor section 24, acombustor section 26 and a turbine section 28. The fan section 22 drivesair along a bypass flow path B in a bypass duct defined within a nacelle15, and also drives air along a core flow path C for compression andcommunication into the combustor section 26 then expansion through theturbine section 28. Although depicted as a two-spool turbofan gasturbine engine in the disclosed non-limiting embodiment, it should beunderstood that the concepts described herein are not limited to usewith two-spool turbofans as the teachings may be applied to other typesof turbine engines including three-spool architectures.

The exemplary engine 20 generally includes a low speed spool 30 and ahigh speed spool 32 mounted for rotation about an engine centrallongitudinal axis A relative to an engine static structure 36 viaseveral bearing systems 38. It should be understood that various bearingsystems 38 at various locations may alternatively or additionally beprovided, and the location of bearing systems 38 may be varied asappropriate to the application.

The low speed spool 30 generally includes an inner shaft 40 thatinterconnects a fan 42, a first (or low) pressure compressor 44 and afirst (or low) pressure turbine 46. The inner shaft 40 is connected tothe fan 42 through a speed change mechanism, which in exemplary gasturbine engine 20 is illustrated as a geared architecture 48 to drivethe fan 42 at a lower speed than the low speed spool 30. The high speedspool 32 includes an outer shaft 50 that interconnects a second (orhigh) pressure compressor 52 and a second (or high) pressure turbine 54.A combustor 56 is arranged in exemplary gas turbine 20 between the highpressure compressor 52 and the high pressure turbine 54. A mid-turbineframe 57 of the engine static structure 36 is arranged generally betweenthe high pressure turbine 54 and the low pressure turbine 46. Themid-turbine frame 57 further supports bearing systems 38 in the turbinesection 28. The inner shaft 40 and the outer shaft 50 are concentric androtate via bearing systems 38 about the engine central longitudinal axisA which is collinear with their longitudinal axes.

The core airflow is compressed by the low pressure compressor 44 thenthe high pressure compressor 52, mixed and burned with fuel in thecombustor 56, then expanded over the high pressure turbine 54 and lowpressure turbine 46. The mid-turbine frame 57 includes airfoils 59 whichare in the core airflow path C. The turbines 46, 54 rotationally drivethe respective low speed spool 30 and high speed spool 32 in response tothe expansion. It will be appreciated that each of the positions of thefan section 22, compressor section 24, combustor section 26, turbinesection 28, and fan drive gear system 48 may be varied. For example,gear system 48 may be located aft of combustor section 26 or even aft ofturbine section 28, and fan section 22 may be positioned forward or aftof the location of gear system 48.

The engine 20 in one example is a high-bypass geared aircraft engine. Ina further example, the engine 20 bypass ratio is greater than about six(6), with an example embodiment being greater than about ten (10), thegeared architecture 48 is an epicyclic gear train, such as a planetarygear system or other gear system, with a gear reduction ratio of greaterthan about 2.3 and the low pressure turbine 46 has a pressure ratio thatis greater than about five. In one disclosed embodiment, the engine 20bypass ratio is greater than about ten (10:1), the fan diameter issignificantly larger than that of the low pressure compressor 44, andthe low pressure turbine 46 has a pressure ratio that is greater thanabout five 5:1. Low pressure turbine 46 pressure ratio is pressuremeasured prior to inlet of low pressure turbine 46 as related to thepressure at the outlet of the low pressure turbine 46 prior to anexhaust nozzle. The geared architecture 48 may be an epicycle geartrain, such as a planetary gear system or other gear system, with a gearreduction ratio of greater than about 2.3:1. It should be understood,however, that the above parameters are only exemplary of one embodimentof a geared architecture engine and that the present invention isapplicable to other gas turbine engines including direct driveturbofans.

A significant amount of thrust is provided by the bypass flow B due tothe high bypass ratio. The fan section 22 of the engine 20 is designedfor a particular flight condition—typically cruise at about 0.8 Mach andabout 35,000 feet (10,668 meters). The flight condition of 0.8 Mach and35,000 ft (10,668 meters), with the engine at its best fuelconsumption—also known as “bucket cruise Thrust Specific FuelConsumption (‘TSFCT’)”— is the industry standard parameter of 1 bm offuel being burned divided by 1 bf of thrust the engine produces at thatminimum point. “Low fan pressure ratio” is the pressure ratio across thefan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The lowfan pressure ratio as disclosed herein according to one non-limitingembodiment is less than about 1.45. “Low corrected fan tip speed” is theactual fan tip speed in ft/sec divided by an industry standardtemperature correction of [(Tram ° R)/(518.7° R)]^(0.5). The “Lowcorrected fan tip speed” as disclosed herein according to onenon-limiting embodiment is less than about 1150 ft/second (350.5meters/second).

FIG. 2 shows an intercooled cooling air system 100. A high pressurecompressor 102 is provided with a tap 104 for tapping air to be utilizedas cooling air. Note the tap 104 is from an intermediate location in thehigh pressure compressor. In alternative embodiments, the tap could befrom a low pressure compressor.

Stated broadly, the tap in the FIG. 2 embodiment is from a locationupstream of a downstream most location 105 in the high pressurecompressor 102. The tapped air at line 104 passes through a first leg106 of a heat exchanger and back into a second leg 108 where it isreturned through an inner housing 109. The heat exchanger 106/108 sitsin a chamber 110 in this embodiment, which is radially inward of thebypass duct B.

As shown, a valve 112 selectively allows the air from bypass duct B topass over the heat exchanger 106/108. A control 114 is shown controllingthe position of the valve 112. Air at line 116 inward of the housing 109passes through a shutoff valve 118 into a cooling compressor 120. Thecooling compressor 120 is provided with a drive 121.

A system to stop the cooling compressor 120 from compressing isprovided. In the illustrated example, the system is a clutch 122 whichcan disconnect the cooling compressor 120 from its drive 121.Alternatively, the drive 121 could be an electric motor and simplystopped. The control 114 controls the shutoff valve 118 and the clutch122. It is desirable that the control be programmed such that thecompression of air by the cooling compressor 120 is effectively stoppedbefore the valve 118 is shut down to block the flow of air from line 116reaching the cooling compressor 120.

Downstream of the cooling compressor 120, the air passes into a line 124and through struts 128 in a diffuser 126 radially into a mixing chamber130. In this embodiment, high pressure air may be tapped at 132 into themixing chamber 130. The high pressure air tapped 132 may be airdownstream of the downstream most location 105 in the high pressurecompressor 102. The air from the mixing chamber 130 is shown passing tocool a disk and rim 142 of a downstream most location in the highpressure compressor 102. The air is also shown passing between a fixedhousing 136 and a rotating shaft portion 140, which is part of the highpressure spool as described with regard to FIG. 1 . A chamber 141between the housing 136 and outer periphery of the shaft portion 140receives the cooling airflow. That air passes through a tangentialon-board injector 144 (“TOBI”) and then passes to cool the first stageblade 146 and vane 148 of a high pressure turbine. The shaft portion 140could be thought of as a rotating portion, and the housing portion 136could be thought of as a non-rotating portion.

As shown, the wall 136 is radially inward of a combustor 134 and achamber 138 is intermediate to the two.

As known, the chamber 138 receives air downstream of the downstream mostlocation in the high pressure compressor and, thus, is at hightemperature. As the air passes to the TOBI 144 and through the chamber141, it may be heated by those high temperatures, which reduces theefficiency of the overall system.

In embodiments, an insulation feature is placed both on at least amajority of the surface area of the housing 136 between the diffuser 126and the TOBI 144. An insulation material is also place along themajority of the outer surface of the shaft portion 140 between thedownstream most location of the high pressure compressor and therotation of the first turbine blade 146.

For purposes of this application, each of the housing 136 and shaft 140are formed of an underlying base metal and an outer insulation material.The outer insulation material has better resistance to heat passage thandoes the underlying metal. Insulation in a gas turbine, and as may bedefined in this application is a non-structural addition to a structurein that there is little or no structural contribution to the additionalaspect involved. Further, the insulation in a gas turbine, may preventfluid from passing on one side of the lower conductivity material whichis sometimes referred to as “infiltration.” Such fluid passage woulddramatically lower the value of the measures take to apply theinsulation. And, finally, the insulation may be introduced to theassembly by additionally manufacturing step and processes.

In an alternative embodiment shown schematically in FIG. 3 , the coolingsystem 200 taps air from a location at or downstream of the downstreammost location 204 in the high pressure compressor 202. Here, the air isshown tapped at 206 from a chamber radially outward of the combustor208. That air passes through a heat exchanger 210, which is shownschematically being cooled, and then returned back inwardly throughchamber 212 to the high pressure turbine 214. Such passage of air mayalso include the wall 136 and shaft 140 as in the FIG. 2 embodiment.Thus, the same temperature challenges are raised.

FIG. 4 shows details of one embodiment of insulation material. The outersurface of the shaft 140, which faces the wall 136, is preferablyprovided with an insulation coating. Coating is preferred for therotating structure as the attachment of sheets or other structure whichmight require mechanical attachment raises challenges with thecentrifugal forces that the rotating structure will see. In thisembodiment, the underlying base metal 150 is provided with a metallicbond coat 152. A thermally-grown oxide coating 154 is placed outwardlyof the bond coat 152. A ceramic topcoat 156 then surrounds thethermally-grown oxide.

In one embodiment, the ceramic topcoat may be composed ofYttria-stabilized zirconia.

An outer surface of the housing 136 includes the metallic base layer 160and a double wall structure, such as provided by an attached outer wall164, which faces the rotating shaft surface 140. An intermediateinsulation material such as a ceramic fiber blanket 162 is placedbetween the walls 160 and 164. Such ceramic fiber blankets are known forvarious applications and may be formed of bulk fibers produced byspinning processes. The blanket 162 may be formed from purealumina-silica. Further, the blanket 162 may be a continuous blanket andmay be mechanically sewn with double needles to provide better integrityto the surface on both sides of the blanket.

In embodiments, the pipes and, particularly, those downstream of theheat exchanger 106/108 (or 210) may also be provided with insulation.This would include connections 116 and, in particular, connection 124.

In addition, monitors are provided to ensure proper operation of valve112, valve 118, clutch 122, and predetermine any undesirable pressuresor temperatures in the conduit 124.

Stated another way, a gas turbine engine include rotatable componentsincluding components within a compressor section and a turbine sectionhoused within an outer housing. A tap is connected to tap air that haspassed at least partially through the compressor section The tap isconnected to pass through a heat exchanger and connected to pass into aflow path between a rotating surface and a non-rotating surface. Theflow path is connected to cool at least one of said rotatablecomponents. At least a portion of each of the non-rotating surface andthe rotating surface are provided with a base metal, and an insulationmaterial on a surface facing the other of the rotatable andnon-rotatable surfaces.

Although an embodiment of this invention has been disclosed, a worker ofordinary skill in this art would recognize that certain modificationswould come within the scope of this disclosure. For that reason, thefollowing claims should be studied to determine the true scope andcontent of this disclosure.

1. A gas turbine engine comprising: a plurality of rotating componentshoused within a compressor section and a turbine section; a first tapconnected to said compressor section and configured to deliver air at afirst pressure; a heat exchanger connected downstream of said first tap;a rotating surface radially outward of a non-rotating surface defining aflowpath therebetween and wherein the non-rotating surface faces therotating surface, wherein the flowpath is connected downstream of saidheat exchanger and is configured to deliver air to at least one of saidplurality of rotating components, wherein at least a portion of saidnon-rotating surface and said rotating surface comprising a base metal;an insulation material disposed on a surface along the flowpath; whereinsaid insulation material being provided outwardly of said base metal onat least a portion of said non-rotating surface; and wherein saidinsulation material on said non-rotating surface includes a ceramicfiber blanket formed of bulk fibers, and wherein said base metal isradially inward of said ceramic fiber blanket and an outer wall of saidnon-rotating structure is attached on an opposed radial side of saidceramic fiber blanket relative to said base metal.
 2. The gas turbineengine as set forth in claim 1, wherein there is a downstream mostlocation in a high pressure compressor within said compressor sectionand the first tap is at an upstream location relative to the downstreammost location.
 3. The gas turbine engine as set forth in claim 1,wherein there is a high pressure compressor in the compressor sectionwith a downstream most location and said first tap is at a locationwhere air will have passed downstream of the downstream most location.4. The gas turbine engine as set forth in claim 1, wherein saidinsulation material is also on said rotating surface.
 5. The gas turbineengine as set forth in claim 4, wherein said insulation of said rotatingsurface is a coating.
 6. The gas turbine engine as set forth in claim 5,wherein said coating including an outer ceramic topcoat facing theinsulation material on said non-rotating surface.
 7. The gas turbineengine as set forth in claim 6, wherein there is a metallic bond coatintermediate said ceramic topcoat and the underlying base metal on saidrotating surface.
 8. The gas turbine engine as set forth in claim 7,wherein there is a thermally-grown oxide coating intermediate saidmetallic bond coat and said ceramic topcoat.
 9. The gas turbine engineas set forth in claim 1, wherein fluid conduits are connected to alocation downstream of said heat exchanger, to communicate airdownstream of the heat exchanger into said flow path, and at least someof said fluid conduits being provided with insulation.
 10. The gasturbine engine as set forth in claim 1, wherein said ceramic fiberblanket is formed with a alumina-silica fibers.
 11. The gas turbineengine as set forth in claim 10, wherein fluid conduits are connected toa location downstream of said heat exchanger, to communicate airdownstream of the heat exchanger into said flow path, and at least someof said fluid conduits being provided with insulation.
 12. The gasturbine engine as set forth in claim 1, wherein there is a combustorradially outward of said non-rotating surface, and a chamberintermediate said combustor and said non-rotating surface connected toreceive compressed air downstream of a downstream most location in saidcompressor section.
 13. A gas turbine engine comprising: a plurality ofrotating components housed within a compressor section and a turbinesection; a first tap connected to said compressor section and configuredto deliver air at a first pressure; a heat exchanger connecteddownstream of said first tap; a flowpath defined between a rotatingsurface and a non-rotating surface, wherein the flowpath is connecteddownstream of said heat exchanger and is configured to deliver air to atleast one of said plurality of rotating components, wherein at least aportion of said non-rotating surface and said rotating surfacecomprising a base metal; an insulation material disposed on a surfacealong the flowpath; wherein said insulation material being providedoutwardly of said base metal on at least a portion of both said rotatingsurface and said non-rotating surface; wherein there is a combustorradially outward of said non-rotating surface, and a chamberintermediate said combustor and said non-rotating surface connected toreceive compressed air downstream of a downstream most location in saidcompressor section; wherein said insulation material on said rotatingsurface is a coating; and wherein said rotating surface is an outersurface of a shaft connecting a high pressure turbine rotor in saidturbine section to a high pressure compressor rotor in said compressorsection.
 14. The gas turbine engine as set forth in claim 13, whereinthere is a downstream most location in a high pressure compressor withinsaid compressor section and the first tap is at an upstream locationrelative to the downstream most location.
 15. The gas turbine engine asset forth in claim 13, wherein there is a high pressure compressor inthe compressor section with a downstream most location and said firsttap is at a location where air will have passed downstream of thedownstream most location.
 16. The gas turbine engine as set forth inclaim 13, wherein said coating including an outer ceramic topcoat facingthe insulation material on said non-rotating surface.
 17. The gasturbine engine as set forth in claim 16, wherein there is a metallicbond coat intermediate said ceramic topcoat and the underlying basemetal in said rotating surface.
 18. The gas turbine engine as set forthin claim 17, wherein there is a thermally-grown oxide coatingintermediate said metallic bond coat and said ceramic topcoat.
 19. Thegas turbine engine as set forth in claim 13, wherein fluid conduits areconnected to a location downstream of said heat exchanger, tocommunicate air downstream of the heat exchanger into said flow path,and at least some of said fluid conduits being provided with insulation.20. The gas turbine engine as set forth in claim 13, wherein said atleast one rotating component includes at least an upstream most bladeand vane in a high pressure turbine which is part of said turbinesection.